Ramjet engine

Received 24 January 2006, revised 19 June 2007

Defence Science Journal, Vol. 58, No. 3, May 2008, pp. 327-337Ó 2008, DESIDOC

327

Keywords: Ramjet engine, air intake, combustor, numerical simulation, internal flow, air/fuel ratio

1 . INTRODUCTION

Ramjet engines are lighter than rockets sincethey utilise ambient air as the oxidiser for combustionof fuel. The major constituents of a liquid ramjetengine are air-intake, combustion chamber, andnozzle. The full engine simulation is useful forunderstanding the various important aspects of ramjetinternal flows that include shockwave/boundarylayer interactions, inlet/combustor coupling, flameholding and spreading, and combustion dynamics.Thomas1 presented documented facts about thecurrent status of ramjets and ramjet technologyand concluded that ramjets possess the uniqueability to provide continuous thrust, sustained highsupersonic speed and high specific impulse. Hebrard2,et al. employed a combined approach using experimentsin isothermal conditions and simple computationmodels to study the global performances of various

Numerical Analysis of Integrated Liquid Ramjet Engine

G. Raja Singh Thangadurai1, B.S. Subhash Chandran1, V. Babu2, and T. Sundararajan2

1Defence Research & Development Laboratory, Hyderabad-500 0582Indian Institute of Technology Madras, Chennai-600 036

ABSTRACT

The numerical simulation of an integrated, liquid-fuelled ramjet engine comprising supersonicair intake, subsonic combustor and a convergent-divergent nozzle has been carried out and theresults are discussed in this paper. These results include cold flow studies, heat addition in thecombustor and full engine analysis with coupled simulation of supersonic air-intake andcombustion chamber along with the nozzle. Overall ramjet operation depends on the performanceof the air intake and the combustion chamber. The coupling phenomena are very dominant andperformance of air intake is affected vastly by the combustor operation and vice versa. In thispaper, a numerical analysis of integrated liquid ramjet engine considering coupling phenomenabetween various sub-systems viz., air intake, combustor and nozzle has been reported.

ramjets. Calzone3 carried out a study of the internationaldevelopments on missile ramjet propulsion. Thechoice and techniques to be used during the designof a liquid-fuelled ramjet engine were described indetail by Cazin and Laurent 4.The specific featuresof the constituents of a ramjet engine such as theair supply system, fuel supply system, combustionchamber and the nozzle, were discussed. Conwayand Johansson5 dealt with the development andvalidation of a methodology for simulation oframjet -powered missiles using computational fluiddynamics (CFD). Simultaneous calculations of theinternal and external flows were carried out bycoupling a quasi one- dimensional ramjet enginemodel with a CFD solver. A thermally perfect gasassumption was adopted to enable accurate modellingof the hot nozzle gas flow. Tip to tail calculationswere carried out on two ramjet-powered vehicles

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DEF SCI J, VOL. 58, NO. 3, MAY 2008

and results were compared with available and windtunnel data. Sung6, et al. numerically analysed anentire ramjet engine to understand the combustiondynamic characteristics of an integrated rocketramjet system.Various physical processes wereinvestigated systematically, including flame propagation,flame dynamics, turbulent flame evolution, terminalshock train, modal analysis of the system, andinteraction mechanism between the oscillatingshockwave in the diffuser and pressure wave resultingfrom heat release, using spectral analysis. Gaiddonand Knight7 used an automated optimisation loopbased on CFD tools for improving the flight performanceof a ramjet-powered missile. For each point of themission (acceleration, cruise, and manoeuvre) thebest inlet shape was determined. Multi-objectiveoptimisations were then performed to obtain thebest set of shapes satisfying the whole mission.The coupling phenomena between the various subsystemsof a liquid ramjet engine, namely the air-intake,combustor, and nozzle have been investigated inthe present study.

2. SIMULATION OF A RAMJET ENGINE

The performance of a ramjet engine dependson the efficient operation of both the air intake andcombustion chamber. Since the operations of thesesubsystems are inter-dependent, there is a requirementto carry out the simulation of a complete ramjetengine in totality to analyse its realistic performancecharacteristics. Complete simulation also helps inchoosing the proper location and size of air inletsas well as the combustor geometry to obtain thebest overall engine performance. However, a veryimportant factor that needs deep consideration isthe interaction between the intake, combustor, andthe nozzle flows. For example, in sub-critical condition,the unattached shock may lead to air spillage, whichcould result in poor combustion and loss of thrust.On the other hand, in super-critical condition duringwhich the shock train moves into the diffuser, pressurerecovery may be poor, again resulting in lowerpropulsion efficiency. The locations of the shockas well as the shock reflections are affected notonly by the geometries of the intake but also bythe amount of heat release in the combustor. Theextent of flow acceleration in the nozzle depends

on both the stagnation pressure and stagnationtemperature conditions achieved in the combustorsection. Thus, a comprehensive understanding ofair intake-combustor-nozzle interactions requires asolution obtained through integrated ramjet simulation.

The full engine simulation is useful in understandingthe various important aspects of ramjet internalflows that include shockwave/boundary layerinteractions, inlet/combustor coupling, flame holdingand spreading, and combustion dynamics. Sincethe focus of this study is on the coupling phenomenabetween the various subsystems, a simple axi-symmetricengine geometry (consisting of axi-symmetric intake,axi-symmetric combustor and axi-symmetric nozzle)has been considered (Fig. 1). In order to achieveaerodynamic flame holding in the combustor, a suddenexpansion has been introduced at the end of theintake. Fuel is injected close to the re-circulationzone, which occurs after the sudden expansion.

In order to clearly highlight the flow phenomenadue to geometry and the effects of spray mixingas well heat release by combustion, the simulationof the complete ramjet engine has been carried outin the following three stages:

Figure 1. Ramjet engine with axi-symmetric intake,combustor and nozzle.

(i) Simulation with cold air flow,

(ii) Simulation with equivalent heat addition at thecombustor wall, and

(iii) Simulation with fuel spray and combustion inthe combustor.

The main parameters which are varied in thefull simulation studies are the flight Mach numberand the air/fuel ratio.

3 . COLD FLOW STUDIES IN A RAMJETENGINE

To simulate the entire ramjet engine, a cylindricalcombustion chamber with a convergent divergent

,

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THANGADURAI, et al.: NUMERICAL ANALYSIS OF INTEGRATED LIQUID RAMJET ENGINE

Figure 2. Grid for integrated engine. Figure 4. Mach contours for cold air flow.

nozzle is attached to the exit section of the airintake. The grid employed for simulation studies isshown in Fig. 2. This arrangement is useful inanalysing the ramjet engine in totality. For thenumerical calculations the system of governingdifferential equations for a steady, compressibleand turbulent flow have been solved using thecommercial software FLUENT with appropriateboundary conditions. On the left hand side (inlet)boundary and on the top and bottom boundaries,the free stream temperature, pressure and Machnumber conditions have been imposed. On the wallsof the engine and on the centre body, adiabaticconditions along with standard wall functions havebeen prescribed. On the right boundary, supersonicoutlet conditions have been prescribed for bothinternal and external flow streams. All the flowvariables including pressure are determined fromthe interior domain by extrapolation.

The turbulence model used in the present simulationstudy is the standard k-e model which employs twopartial differential equations to estimate the velocityand length scales of turbulence. The standard wallfunctions described by Launder and Spalding8 areused in the present calculations. Numerical simulationswere carried out using three different grid sizeswith 100 x 100, 150 x 150, and 200 x 200 nodesin the axial and radial directions to obtain solutionsfor cold air flow through the engine. Based on gridsensitivity analysis, grid size of 200 x 200 has beenchosen for all calculations. The predicted staticpressure and Mach contours for cold flow analysisof a typical case are shown in Figs 3 and 4 respectively.With the enlarged combustion chamber the terminalnormal shock moves further downstream inside

the air duct. The oblique shocks attached to thefore body, tip of the cowl and also near the suddenexpansion section in the external flow, are clearlyseen in the static pressure and Mach contours. Itis interesting to note that the peak pressure valueis achieved only just before the nozzle section.

At a higher Mach number, the static pressurerecovery is also greater as expected. A closerscrutiny of the Mach number contours revealsreflected oblique shock patterns in the intake region.Expansion of the flow (compressed by shock inthe intake) is observed across the nozzle.Stream

Figure 3. Static pressure (bar) contours for cold air flow.

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function contours are shown in Fig. 5. A re-circulationzone is observed where there is sudden enlargementin the combustion chamber area. The size of there-circulation eddy is seen to increase with Machnumber.

4 . SIMULATION WITH HEAT ADDITIONIN THE COMBUSTOR

For the simulation of hot flow in a ramjetengine, the thermal input for a typical ramjet enginewas distributed over the combustion chamber wallas a uniform heat flux. Steady state combustion ofthe combustible mixture in a ramjet combustor fora fixed set of flow variables generates a constantheat input. Uniform heat flux distributed over thecylindrical combustion chamber wall produces aconstant heat input. The heat input is arrived at byconsidering the following nominal conditions thatprevail in a typical practical ramjet engine: Air/fuelratio = 15; air mass flow rate = 7 kg/s; fuel flowrate = 0.4667 kg/s. Kerosene (JP-8) is used as thefuel for simulation studies. For the above conditionsthe heat input works out to be 10 MW.

The parametric study has been carried out forheat inputs of 10 MW, 12 MW and 15 MW. As inthe case of cold flow simulation, grid size of200 x 200 has been chosen for carrying out calculationswith heat addition in the combustion chamber. Fora given geometry, at a fixed free stream Machnumber and altitude of operation, the mass flowrate through the intake is constant. This mass flowrate (for a given throat area) is proportional to p

o/

To

1/2. When the stagnation temperature is increaseddue to heat addition in the combustion chamber, themass flow rate through the combustor has to bedecreased. To achieve this, the normal shock movesupstream where the Mach number is higher. Thisresults in increased loss of stagnation pressurethereby reducing the mass flow rate. With heataddition, the terminal normal shock moves upstreamand is located near the narrowest cross section ofthe air intake. The contours of static pressure andMach number for this condition shown in Figs 6and 7 corroborate this observation. Due to theforward movement of the terminal shock, the peakpressure value is attained in the rear part of the

intake zone itself, for free stream Mach numbersof 2.0 and 2.5. For a free stream Mach numberof 3.0 also, most of the pressure recovery occursin the intake, although a small pressure rise is seenin the combustor zone.

However, a closer scrutiny of the pressurevalues indicated a small drop (Fig. 6) in the rearpart of the combustor, which can be attributed tothe effect of heat addition in the constant areasection. Thus, the Rayleigh flow in the constantarea combustion chamber also has been predictedwell and the pressure drop across the combustormatched well with the predictions of the corresponding

Figure 5. Stream function (m3/s) contours for cold air flow.

Figure 6. Static pressure (bar) contours for air flow with heataddition (12 MW) in the combustor.

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one-dimensional analysis. The peak pressure valuesattained for the cases with heat addition are higherthan the corresponding values shown for cold flowcase (Fig. 3), for each Mach number considered.Thus, it is evident that the shock locations and theshock compression process are intimately coupledwith the heat addition due to combustion in thecombustion chamber. The Mach number contoursdepicted in Fig. 7 illustrates that the terminal normalshock (corresponding to the transition from M>1to M<1) is located at the entry section for the freestream Mach number of 2.0, while the terminalshock location moves to the constant area sectionof the intake for M = 2.5 and M = 3.0. Thestreamline contours in the combustion chamberportion are shown in Fig. 8. It is seen that re-circulatory eddies are present near the suddenexpansion for the case with heat addition also.

Moreover, flow separation is observed at theend of the constant area portion of the intake forM = 2.5 and M = 3.0. The static pressure distributionalong the centre body outer surface and axis for thecases, with and without thermal input in the combustionchamber inner wall, are shown in Fig. 9. It is evidentthat heat addition significantly increases the staticpressure recovery for the given flow geometry. However,for a given free stream Mach number as the heataddition is increased, the rate of increase in static

pressure rise becomes less. These trends can beattributed to the changes in the shock locations andtheir strength due to heat addition. In Fig. 10, theeffects of heat addition on the Mach number distributionin the axial direction are shown. It is evident that theflow decelerates from supersonic to subsonic conditionbecause of the terminal normal shock. Beyond thisshock, in the combustor section, slight flow accelerationis seen even before the nozzle portion. For the coldflow case, it is clear that the terminal shock occursvery much inside the combustor region. The axialvariation in static temperature is depicted in Fig. 11for different free stream Mach numbers.

Static temperature increases across the obliqueshock and the terminal normal shock in the intakeregion. The Mach number continues to increase inthe combustor region due to heat addition.The capturedair mass flow rate is different for different Machnumbers and hence temperature rise ( / )pT Q m cD = & &

due to heat addition varies with Mach number.

5 . SIMULATION WITH COMBUSTIONIN THE COMBUSTOR

Simulation studies were also carried out for afull liquid ramjet engine comprising all the constituentassemblies such as airintake, fuel injector, combustionchamber and nozzle. The exact operation has beensimulated by injecting fuel in the combustor, immediately

Figure 7. Mach contours for air flow with heat addition(12 MW) in the combustor.

Figure 8. Stream function (m3/s) contours for air flow withheat addition (12 MW) in the combustor.

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0.0 0.2 0.4 0.6 0.8 1.0

0.0

2.0x105

4.0x105

6.0x105

Free stream Mach number 2.0

MACH 2.0

COLD AIR FLOW

10 MW HEAT ADDITION

12 MW HEAT ADDITION

15 MW HEAT ADDITION

Sta

ticpre

ssure

(bar)

Distance along the axis ( m )

0.0 0.2 0.4 0.6 0.8 1.0

0.0

2.0x105

4.0x105

6.0x105

8.0x105

1.0x106

Free stream Mach number , M = 2.5

COLD AIR FLOW

10 MW HEAT ADDITION

12 MW HEAT ADDITION

15 MW HEAT ADDITION

Sta

ticpre

ssure

(bar)


0.0 0.2 0.4 0.6 0.8 1.0

0.0

2.0x105

4.0x105

6.0x105

8.0x105

1.0x106

1.2x106

Free stream Mach number M = 3.0

COLD AIR FLOW

10 MW HEAT ADDITION

12 MW HEAT ADDITION

15 MW HEAT ADDITION

Sta

ticpre

ssure

(bar)


Figure 9. Effect of heat addition on the static pressuredistribution along the axis.

after the sudden expansion section. The computationaldomain is given in Fig.12

The effect of free stream Mach numberand air/fuel ratio (A/F) on the ramjet engineperformance have been studied in detail. Full

engine simulations have been carried out withthree different grids having 11090, 22080 and44350 cells. Refinement of the grid has beencarried out at the location where steep gradientsof temperature and pressure are seen. All the

0.0 0.2 0.4 0.6 0.8 1.0

0.0

0.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0


COLD FLOW

10 MW HEAT ADDITION

12 MW HEAT ADDITION)

15 MW HEAT ADDITION

Mach

num

ber


0.0 0.2 0.4 0.6 0.8 1.0

0.0

0.5

1.0

1.5

2.0

2.5

3.0


COLD FLOW

10 MW HEAT ADDITION

12 MW HEAT ADDITION

15 MW HEAT ADDITION

Mach

num

ber


0.0 0.2 0.4 0.6 0.8 1.0

0.0

0.5

1.0

1.5

2.0

2.5


COLD FLOW

10 MW HEAT ADDITION

12 MW HEAT ADDITION

15 MW HEAT ADDITION

Mach

num

ber


Figure 10. Effect of heat addition on the Mach number distribution along the axis.

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0.0 0.2 0.4 0.6 0.8 1.0

300

350

400

450

500

550

Free stream Mach number, M = 2.0

COLD FLOW

10 MW HEAT ADDITION

12 MW HEAT ADDITION

15 MW HEAT ADDITION


0.0 0.2 0.4 0.6 0.8 1.0

300

350

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450

500

550

600

650

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COLD FLOW

10 MW HEAT ADDITION

12 MW HEAT ADDITION

15 MW HEAT ADDITION


0.0 0.2 0.4 0.6 0.8 1.0200

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COLD FLOW

10 MW HEAT ADDITION

12 MW HEAT ADDITION

15 MW HEAT ADDITION


Sta

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Te

mp

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ture

(K)

Sta

tic

Te

mp

era

ture

(K)

Sta

tic

Te

mp

era

ture

(K)

COMBUSTIONCHAMBERAIR INTAKE

CENTRE BODY

FUEL SPRAY NOZZLE

Figure 12. Computational domain.

Figure 11. Effect of heat addition on the static temperaturedistribution along the axis.

and 3.0 are shown in Figs 13 and 14 respectively.The ir/fuel ratio for this case was 17. Mach numbercontours for inlet Mach number equal to 2 (Fig. 13)indicate that the shock train is located outside theintake, which results in spillage of air flow. For M

i

= 2.5, the spillage is reduced and the terminal normalshock is located at the entry section. For M

i = 3.0,

the shock train moves into the intake; here, thefeature of reflected oblique shocks culminating ina terminal shock is clearly seen.These three casesof M

i = 2.0, 2.5 and 3.0 correspond to the sub-

critical, critical (approximately) and super-criticaloperation of the engine respectively. From the temperaturecontours (Fig. 14) it is seen that combustion primarilyoccurs close to the wall, near the sudden expansionsection.The re-circulatory region and boundary layerclose to the wall aid in flame stabilisation. In thisstudy, the conserved-scalar approach based on thefast-chemistry assumption is adopted to accountfor the turbulence-combustion interaction. The maximumtemperature attained is highest for M

i = 3.0, since

the shock compression process results in a higherpre-ignition temperature for this case because ofstronger shocks.

Figure 13. Mach contours for air/fuel ratio = 17.

results reported here have been obtained on thefinest grid, i.e., with 44350 cells.

5.1 Effect of Free Stream Mach Number

The contours of Mach number and statictemperature for inlet Mach numbers of 2.0, 2.5

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Although maximum temperature is higher, theaverage temperature is lower for M

i = 3.0 due to

larger mass flow rate. In Fig. 15, the fuel massfraction contours are plotted. It is clear that fueldisperses more at M

i = 2.0 as compared to the

cases of Mi = 2.5 and 3.0. Therefore, the combustion

zone is larger for Mi = 2.0. The static pressure,

static temperature and Mach number variation alongthe surface of the centre body and axis are shownin Figs 16, 17 and 18 respectively. The static pressureand temperature increase across the terminal normal

Figure 15. Mass fraction contours of fuel for air/fuelratio = 17.

Figure 14. Temperature (K) contours for air/fuel ratio.

shock, while the Mach number decreases to subsonicvalues. In fact, for M

i = 3.0, features such as flow

deceleration at the first shock, re-accelerationimmediately after the shock and subsequent transitionto subsonic flow at the terminal normal shock canbe clearly discerned. Also, at M

i = 3.0, combustion

phenomenon does not penetrate up to the axis andhence temperature rise is marginal along the axis.

5.2 Effect of Air/Fuel Ratio

Results for an air fuel ratio of 30 are presentedin Figs. 19 and 20. Compared with the previouscase for which the air/fuel ratio was 17, significantchanges can now be seen. For example, the Machnumber contours show that the ramjet operation

0.0 0.2 0.4 0.6 0.8 1.0

0.0

2.0x105

4.0x105

6.0x105

8.0x105

1.0x106

1.2x106

AIR / FUEL RATIO = 17

Mi = 2.0

Mi = 2.5

Mi = 3.0

DISTANCE ALONG THE COMBUSTOR AXIS ( M )

STA

TIC

PR

ES

SU

RE

(Pa)

Figure 16. Variation of static pressure alng the surface andaxis of the centre body.

0.0 0.2 0.4 0.6 0.8 1.0

200

400

600

800

1000

1200

1400


DISTANCE ALONG THE COMBUSTOR AXIS (M)

M i = 2.0

M i = 2.5

M i = 3.0

STA

TIC

TE

MP

ER

AT

UR

E(K

)

Figure 17. Variation of temperature along the surface andaxis of the centre body.

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Figure 18. Mach number variation along the surface and axisof the centre body.

Figure 20. Temperature (K) contours for air/fuel ratio= 30.

an A/F of 30. As a result of changes in the flowstructure and air flow rate, combustion is poorernear the axis for A/F = 30. Thus, it is evident thatthe air fuel ratio plays a crucial role for theproper operation of a ramjet engine.

6 . CONCLUSIONS

The overall performance of an integrated ramjetengine comprising of air-intake, combustor andnozzle has been investigated through full enginesimulation. The coupled analysis clearly brings

becomes super-critical (shocks moves within theintake) even for M

i = 2.5. Also, the extent of spillage

decreases for Mi = 2.0. In addition, the inward movement

of the shock system, results in higher temperaturevalues for M

i = 2.5 and 3.0. However, most of the

features of the combustion zone such as fuel dispersion,(Fig. 21) are similar to the corresponding predictionsfor A/F = 17. In Figs 22 - 24, the centre line variationsof static pressure, static temperature and Mach numberalso indicate the slight shift in the location of thenormal shock at M

i = 2.5. Consequently, the flow

and pressure recovery are modified. For inlet Machnumber of 2.5, the peak temperature observed foran A/F of 17 is 1980 K as compared to 1920 K for

Figure 19. Mach contours for air/fuel ratio= 30.

Figure 21. Mass fraction contours of fuel for air/fuelratio= 30.

0.0 0.2 0.4 0.6 0.8 1.0

0

1

2

3

4


Mi = 2.0

M i = 2.5

M i = 3.0


MA

CH

NU

MB

ER

m

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out the interaction between the air-intake and combustor.The air-intake performance is not only affected byits geometry but also by the amount of heat releasein the combustor. The flow model developed in thepresent study has the potential to be used as adesign and development tool in liquid fuelled ramjetdevelopment programmes.

ACKNOWLEDGEMENTS

The authors are extremely thankful to Shri P.Venugopalan, Director; and Dr S. Sundarrajan, Head,Programme Team, PJ-10, Defence Research andDevelopment Laboratory (DRDL), Hyderabad, fortheir constant encouragement to complete this worksuccessfully.

REFERENCES

1. Thomas, Jr. & A.N. Exploding ramjet myths.National Defense, 1983, 18-23.

2. Hebrard, P; Lavergne, G. & Torque, A. Ramjetand ramrocket performances: Experimentalsimulation and prediction by computation codes.In 10th ISOABE Proceedings, UK, 1991. pp.1051-059. 1991. ISABE 91 - 7112.

3. Calzone ,R.F. Developments in missile Ramjetpropulsion. TNO report. PML 1996 - A100.1991.

4. Cazin, P. & Laurent, J.M. Liquid-fuelled ramjetengine tactical missile propulsion. Prog. Astro.Aero., 1996, 170, 423-46.

5. Conway, S. & Johansson, J. Integrated engine-airframe calculations using CFD. 2001. AIAAPaper No. 2001-3201.

6. Sung, H.G.; Hsieh, S.Y. & Yang, V. A unifiedanalysis of Ramjet operation in an integratedrocket ramjet engine. Part II Combustion dynamicsof ramjet engine. AIAA Paper No. 2001-3192.

7. Gaiddon, A. & Knight, D.D. Aerodynamicoptimization of the aeropropulsive system of aramjet powered missile. 2002. AIAA PaperNo. 2002-5546.

8. Launder, B.E. & Spalding, D.B. The numericalcomputation of turbulent flows. Computer MethodsAppl. Mech. Engg., 1974, 3, 269-89.

0.0 0.2 0.4 0.6 0.8 1.0

200

400

600

800

1000

1200

1400

M i = 2.5




STA

TIC

TE

MP

ER

AT

UR

E(K

)

Figure 23. Variation of static temperature along surface andaxis of the centre.

0.0 0.2 0.4 0.6 0.8 1.0

0.0

2.0x105

4.0x105

6.0x105

8.0x105

1.0x106

1.2x106

Mi = 2.5



STA

TIC

PR

ES

SU

RE

(Pa

)


Figure 22.Variation of static pressure along the surface andaxis of the centre body.

0.0 0.2 0.4 0.6 0.8 1.0

0

1

2

3

M i = 2.5




MA

CH

NU

MB

ER

Figure 24. Mach number variation along the surface andaxis of the centre body.

m

m

m

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9. Sreenatha, A.G. & Bhardwaj, N. Mach numbercontroller for a flight vehicle with ramjet propulsion.1999. AIAA Paper No. 99-2941 IP.

10. Mathew, G.; Lazar, T.C.; Debasis, C.; Paul,P.J. & Mukunda, H.S. Computational studieson the flow field in an ejector ramjet engine.Paper presented at National Conference on Air

Breathing Engines and Aerospace Propulsion,Hyderabad, 1999, pp. 364-74.

11. Minard, J.P.; Hallais, M. & Falempin, F. Lowcost ramjet technology for tactical missileapplication. AIAA Paper No. 2002 - 3765.In 38th AIAA/ASME/SAE/ASEE Joint PropulsionConference & Exhibit, Indianapolis, Indiana.2002.

Contributors

Dr G. Raja Singh Thangadurai obtained his ME (Aero Engg) and PhD (MechEngg) from the Indian Institute of Science (IISc), Bangalore, in 1991 and theIndian Institute of Technology (IIT) Madras, Chennai, in 2004, respectively. Presently,he is working as Scientist E at the Defence Research & Development Laboratory(DRDL), Hyderabad. He is involved in the development of liquid propellant rocketengines and reaction control systems for Prithvi and Agni missiles. His areasof research are rocket and ramjet propulsion, numerical simulation of internalflows, and supersonic air intakes. He has contributed 10 papers in national journalsand conferences.

Dr B.S. Subhash Chandran obtained his PhD (Aero Engg) from the Georgia Tech,USA, in 1984. Presently, he is working as Scientist at the DRDL, Hyderabad. Heis actively engaged in the development of Akash propulsion system and liquidfuel ramjet engine for flight vehicles. His areas of interest are rocket and ramjetpropulsion and supersonic air intakes. He has published 15 papers in nationaland international conferences.

Dr V. Babu obtained his PhD (Mech) from the Ohio State University, USA, in1991. Presently, he is working as Associate Professor at the IIT Madras. His areasof research include: Computational simulation of internal/external flows, simulationof chemically reacting , plasma and nonequilibrium flows, high performance computingand development of software tools for engineering analysis. He is recipient ofHenry Ford Technology Award (1998) presented by the Ford Motor Co, UK. forthe design, development and deployment of a virtual aerodynamic/aero-acousticwind tunnel. He has four patents to his credit.

Prof T. Sundararajan obtained his PhD (Mech Engg) from the University ofPhiladelphia, USA, in 1983. He worked as a postdoctoral fellow at the Univerityof Philadelphia, USA, from 1983-84. He joined as Assistant Professor at the IIT,Kanpur, in 1985. Presently, he is working at the IIT Madras, as Professor. He hasguided 17 students for their PhD and 24 students for their MS. He has published76 research papers in various journals and presented 85 papers in various conferences.He has also published a textbook on computational fluid dynamics. His areas ofresearch include: Spray combustion, jet flows, heat transfer and fluid flow inporous media, and thermal modelling of manufacturing and metallurgical problems.

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